1. Field of the Invention
The present invention relates generally to a gas turbine engine and, more particularly, to an aft entry system and method for supplying cooling air to turbine stages of an aircraft turbine engine.
2. Description of the Prior Art
A gas turbine engine of the turbofan type generally includes a forward fan and booster compressor, a middle core engine, and an aft low pressure power turbine. The core engine encompasses a compressor, a combustor and a high pressure turbine in a serial flow relationship. The compressor and high pressure turbine of the core engine are interconnected by a central shaft. The compressor is rotatably driven to compress air entering the core engine to a relatively high pressure. This high pressure air is then mixed with fuel in the combustor and ignited to form a high energy gas stream. This gas stream flows aft and passes through the high pressure turbine, rotatably driving it and the core engine shaft which, in turn, rotatably drives the compressor.
In the turbofan engine, the residual gas stream leaving the core engine high pressure turbine is expanded through a second turbine, which as mentioned above is the aft low pressure turbine. The aft low pressure turbine, in turn, drives the forward fan via a separate shaft which extends forwardly through the central shaft of the high pressure turbine rotor. Although some of the thrust is produced by the residual gas stream exiting the core engine, most of the thrust produced is generated by the forward fan.
It is common practice with respect to gas turbine engines to provide some form of cooling for the hot regions of the turbine engine. This cooling has mainly involved the use of air bled from the compressor of the engine which is then fed to the regions of the engine to be cooled. Thereafter, the air is allowed to rejoin the main gas flow of the engine.
The efficiency of a gas turbine engine is dependent upon many factors. One factor is the degree to which high pressure air generated by the compressor of the engine and intended primarily for driving the high pressure turbine after passage through the combustor is siphoned or bled off to other uses in the engine. One such use of bleed air is the aforementioned cooling of metal surfaces in the hot regions of the engine to maintain them sufficiently to obtain useful strength properties. The greater the amount of high pressure air diverted to other uses in the engine, the less the amount of air to drive the core turbine and thus the less efficiently the high pressure turbine will operate.
Heretofore, the metal surfaces of the stator and rotor blades composing the high and low pressure turbines of the engine have typically been cooled by a forward air entry cooling flow system. In such forward air entry cooling system, high pressure air from the compressor is used to cool the low pressure turbine. The high pressure air is routed axially under the high pressure turbine located immediately forward of the low pressure turbine.
Significant drawbacks exist with the use of high pressure compressor air for cooling and employment of the forward air entry cooling system. One drawback is the reduction of high pressure turbine efficiency. Another drawback is the difficult and complicated task of passing the high pressure compressor air flow underneath the high pressure turbine in view of the presence of components such as the bearings of the high pressure turbine rotor.
Consequently, a need exists for an alternative system for cooling hot regions of an engine, such as the high and low pressure turbine stages of the gas turbine engines, which will avoid the above-mentioned drawbacks.